The subject matter disclosed herein relates to a blade for a turbine, such as in an aircraft engine, a gas turbine engine, a steam turbine, etc. More specifically, the present invention relates to the cooling of a turbine blade tip shroud.
A gas turbine engine comprises a turbine section wherein hot compressed gas is expanded to produce rotating shaft power. The turbine section often comprises a plurality of alternating rows of stationary vanes (nozzles) and rotating blades (buckets). Each rotating blade has an airfoil and a root that attaches the rotating blade to a rotor.
In some cases, an integral tip shroud is included on the radially outward end of each turbine blade so that, when assembled, a set of blades create an outer surface for constraining the passage of the hot compressed gases through the airfoil sections of the blades. The incorporation of integral tip shrouds tends to increase the ability of a turbine section to extract work from the hot compressed gases, improving performance of the turbine engine. Unfortunately, integral tip shrouds on rotating airfoils are highly stressed due to the mechanical and aerodynamic forces, and the high temperature environment, to which they are subjected.
To improve the useful design life of a turbine blade, cooling methods are employed. Traditionally, blade cooling is accomplished by extracting a portion of the compressed working fluid (e.g., air) from the compressor and passing it directly to the turbine section without exposing the cooling fluid to the addition of heat in the combustor section. This cooling fluid provides a source of pressurized and relatively cool fluid, which readily flows through passages formed in the turbine blades and provides cooling thereto. Thus, radial passages are often provided to carry cooling fluid radially outwardly from a root of the blade to a blade's tip where the cooling fluid is discharged.
Accordingly, those skilled in the art seek a turbine blade with improved cooling for airfoil trailing edge and tip shroud.